The present invention relates to gas turbine nozzles and shrouds in the hot gas path of a turbine and which nozzles and shrouds are preferentially located relative to a circumferential array of combustors based on inlet conditions to the nozzles and shrouds, i.e., known circumferential flow characteristics of the hot combustion gases flowing through the nozzle inlet plane and shroud inlet plane.
In conventional gas turbines designed for industrial use, e.g., generation of electrical power, the combustion system is typically comprised of an annular array of circumferentially spaced combustors. Each combustor provides hot gases of combustion through an associated transition piece for flow over a given span of the first-stage nozzles, a given span of first-stage shrouds opposite the first-stage turbine buckets and then through the nozzles and past shrouds of later stages. With respect to the nozzles, each nozzle is comprised of a pair of circumferentially spaced, adjacent nozzle vanes and inner and outer sidewalls defining the flowpath through the nozzle for the hot gases of combustion. In the design of combustors, there are known variations in the circumferential flow characteristics that cause each nozzle to see different inlet conditions. For example, at the first-stage nozzle inlet plane or substantially at the transition piece exit plane, one nozzle may see significantly different heat transfer coefficients and/or temperatures than an adjacent nozzle receiving hot gases of combustion from the same combustor and transition piece. Moreover, one of the nozzles of the set of nozzles which receives the hot gases of combustion from a single combustor may see different flow conditions at different locations along the nozzle inlet. For example, in a gas turbine combustion system having fourteen combustors and forty-two first-stage nozzles, it will be appreciated that the combustor/nozzle clocking arrangement provides three nozzles which receive the hot gases of combustion from one combustor. Because of the variations in flow characteristics, the inlet conditions seen by one of the nozzles are considerably different from the inlet conditions seen by the other two nozzles. More particularly, because of the swirling effects of the flow of fuel within the combustor, a first nozzle of the three nozzles not only may have a higher temperature buildup than the two adjacent nozzles but also a higher temperature at a location along the outer diameter and adjacent an outer corner of the nozzle. The other two nozzles of the set of nozzles receiving hot gases of combustion from the one combustor may have substantially the same inlet temperature uniformly across each nozzle inlet. A hot spot is thus created in a first-stage nozzle of each set thereof associated with each combustor and which hot spot which can vary in temperature as much as 500xc2x0 F. relative to the remaining nozzles of the set. The different flow characteristics also produce variations in pressure.
Because of those recognized variations in the circumferential flow and temperature characteristics at the inlet plane of nozzles, nozzle components are conventionally uniformly designed to meet the most deleterious combustor conditions. As a consequence, one or more nozzles of each nozzle set will be over-designed, which has a negative effect on engine performance and cost. For example, the first-stage nozzle of an industrial gas turbine is typically air or steam-cooled. By designing all nozzles of a stage the same and for the worst-case scenario, the first nozzle which sees a higher temperature inlet condition than the two adjacent nozzles of the set of nozzles receiving combustion gases from the same combustor may be cooled adequately for that condition. However, the other nozzle(s) of that set will be over-cooled, using valuable compressor discharge air or steam with consequent negative impact on engine performance. Further, nozzles for industrial gas turbines are typically formed in nozzle segments and secured in a circumferential array thereof to form the first and second-stage nozzles. Notwithstanding tight controls on production, each nozzle segment may have a different quality. For example, the welds on the nozzle segments may be different or the magnitude of the thermal barrier coatings may be slightly different. Consequently, the structural characteristics of the segments may have slight variations which may lead to acceptance or non-acceptance of the segments for use in the gas turbine. The structural characteristics of each nozzle segment may therefore be unacceptable for forming a nozzle at a xe2x80x9chot spotxe2x80x9d but perfectly acceptable for a nozzle at a different location within the same set which would be subject to less stringent conditions.
The same is true for the shrouds surrounding the buckets for the various turbine stages. Thus, the shrouds of the various stages similarly see the variations in the circumferential flow characteristics along a shroud inlet plane. The shrouds therefore see significantly different heat transfer coefficients and/or temperatures than adjacent shrouds receiving the hot gases of combustion from the upstream nozzle stage. Similarly as the nozzles, the shrouds are conventionally uniformly designed to meet the most deleterious flowpath conditions with the over-designed shrouds having similar negative effects on engine performance and cost as the nozzles as previously described.
Typically, there are the same number of inner shrouds as nozzles. Alternatively, there may be a different number of shrouds than nozzles, for example, two shrouds for each nozzle. In any event, the various shrouds about the hot gas path will see different inlet conditions as described above.
In accordance with a preferred embodiment of the present invention, the nozzles and shrouds of each set of nozzles and shrouds for each associated combustor are preferentially located in accordance with respective nozzle and shroud inlet conditions. For example, and with respect to the nozzles, where hot spots in the inlet conditions to each set of nozzles receiving the hot gases of combustion from an associated combustor are identified, the nozzle at that circumferential location can be designed for that increased temperature condition. Thus, that nozzle may be provided with increased cooling, e.g., increasing the air or steam flow through the nozzle to further cool the nozzle to accommodate the hot spot. Conversely, the remaining nozzle or nozzles of the set of nozzles receiving the combustion gases from the same combustor need not be designed to the worst-case scenario but can be designed, for example, to provide a reduced cooling flow of air or steam. In this manner, over-design of the latter nozzle(s) is avoided. Also, the quality of the nozzles forming a set of nozzles receiving combustion gases from one combustor can be different. For example, the structural quality of the nozzles receiving the cooler flow of the hot gases of combustion need not have the same structural quality of the nozzle of that set which receives the hotter flow from the same combustor. Thus, wall thicknesses or coatings such as thermal barrier coatings, or both, can be reduced for those nozzles identified with the cooler flows of combustion gases as compared with the wall thickness and/or coatings of the nozzle of that set which receives the hotter gases of combustion gases. By preferentially designing the nozzles of each set thereof and locating those nozzles dependent upon the inlet conditions from each combustor, engine performance and total life of the nozzles can be increased. It will be appreciated that the foregoing is applicable to both first and second-stage nozzles.
Similarly as with the case of the nozzles, the shrouds are preferentially located in accordance with conditions of the hot gases flowing along the hot gas path past an inlet plane to the annular array of shrouds of the rotor stages. For example, where hot spots in the inlet conditions to shrouds downstream of the nozzles are identified, the shroud or shrouds at that location can be designed for that increased temperature condition. For example, increased cooling may be supplied. Conversely, the shrouds of the remaining set of shrouds receiving the combustion gases from the same combustor, albeit via the upstream nozzles, need not be designed to the worst-case scenario but can be designed to provide reduced cooling or reduced structural quality. Moreover, thicknesses or coatings can be altered to accommodate the reduced temperature of the gases. Accordingly, the shrouds, similarly as the nozzles, can be preferentially designed dependent upon conditions of the hot gases of combustion flowing past the shrouds for increased engine performance and total life of the shrouds. It will also be appreciated that the foregoing is applicable to the shrouds of each of the turbine stages.
In a preferred embodiment according to the present invention, there is provided for a gas turbine having a circumferential array of components at least in part defining a hot gas path through the turbine and a plurality of combustors for flowing hot gases of combustion through respective sets of components, first and second components of each set of components being subject to different inlet conditions of the hot gases of combustion from an associated combustor, a method of placement of the components and combustors relative to one another, comprising the step of preferentially locating the first component relative to the second component within each set of components at a circumferential location relative to the associated combustor based on the different inlet conditions to the components.
In a further preferred embodiment according to the present invention, there is provided for a gas turbine having a circumferential array of nozzles and a plurality of combustors for flowing hot gases of combustion through respective sets of adjacent nozzles, first and second nozzles of each set of nozzles being subject to different inlet conditions of the hot gases of combustion from an associated combustor, a method of placement of the nozzles and combustors relative to one another, comprising the step of preferentially locating the first nozzle relative to the second nozzle within each set of nozzles at a circumferential location relative to the associated combustor based on the differential inlet conditions to the nozzles.
In a further preferred embodiment according to the present invention, there is provided for a gas turbine having a circumferential array of components defining at least in part a hot gas path through the turbine and a plurality of combustors for flowing hot gases of combustion through respective sets of components, first and second components of each set of components being subject to different inlet conditions of the hot gases of combustion from an associated combustor, a method of placement of the components and combustors relative to one another, comprising the step of increasing turbine performance by preferentially locating the first component relative to the second component within each set of components at a circumferential location relative to the associated combustor based on the different inlet conditions.
In a further preferred embodiment according to the present invention, there is provided for a gas turbine having a circumferential array of components defining at least in part a hot gas path through the turbine and a plurality of combustors for flowing hot gases of combustion through respective sets of components, first and second components of each set of components being subject to different inlet conditions of the hot gases of combustion from an associated combustor, a method of placement of the components and combustors relative to one another, comprising the step of increasing the part life of the components by preferentially locating the first component relative to the second component within each set of components at a circumferential location relative to the associated combustor based on the different inlet conditions.
In a further preferred embodiment according to the present invention, there is provided a gas turbine comprising a circumferential array of components at least in part defining a hot gas path through the turbine, a circumferential array of combustors for flowing hot gases of combustion along the hot gas path through respective sets of adjacent components, first and second components of the sets thereof being subject to different inlet conditions of the hot gases of combustion from respective combustors associated therewith, the first component of each set thereof being located at a circumferential location relative to the second component of each set thereof and the associated combustor based on the different inlet conditions and having a qualitative difference in comparison with the second component.